This invention relates generally to turbine assemblies, and more particularly, to methods and apparatus for cooling gas turbine engine rotor blade tips.
Gas turbine engines typically include a compressor, a combustor, and at least one turbine. The compressor compresses air which is mixed with fuel and channeled to the combustor. The mixture is then ignited for generating hot combustion gases, and the combustion gases are channeled to the turbine which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator.
The turbine includes a rotor assembly and a stator assembly. The rotor assembly includes a plurality of rotor blades extending radially outward from a disk. More specifically, each rotor blade extends radially between a platform adjacent the disk, to a tip. A combustion gas flowpath through the rotor assembly is bound radially inward by the rotor blade platforms, and radially outward by a plurality of shrouds.
The stator assembly includes a plurality of stator vanes which form a nozzle that directs the combustion gases entering the turbine to the rotor blades. The stator vanes extend radially between a root platform and a tip. The tip includes an outer band that mounts the stator assembly within the engine.
During operation, the turbine stator and rotor assemblies are exposed to hot combustion gases. Over time, continued exposure to hot combustion gases increases an operating temperature of the rotor assembly. As the rotor assembly rotates, higher temperatures migrate from each rotor blade root towards each rotor blade tip. The increased operating temperature of the rotor blade tips may cause the shroud surrounding the rotor assembly to weaken and oxidize.
To facilitate reducing operating temperatures of the rotor blade tips, at least some known rotor assemblies include blade cooling systems which channel cooling air from a compressor through a pre-swirl system. The pre-swirl system discharges the air into radial passages in the rotor blades. The cooling air flows through the rotor blades and is exhausted radially outward through the tip of the blade. Such cooling systems are costly and use significant amounts of cooling air in addressing a local, life-limiting problem.
In an exemplary embodiment, a turbine for a gas turbine engine includes a turbine nozzle assembly that facilitates reducing an operating temperature of rotor blades in a cost-effective and reliable manner. Each rotor blade includes a tip that rotates in close proximity to a shroud extending circumferentially around the rotor assembly. The turbine nozzle assembly includes a plurality of turbine vane segments that channel combustion gases to downstream rotor blades. Each turbine vane segment extends radially outward from an inner platform and includes a tip, a root, and a body extending therebetween. The turbine vane segment tip is formed integrally with an outer band used to mount the vane segments within the gas turbine engine. The outer band is in flow communication with a cooling fluid source, and includes at least one opening.
During operation, as the turbine rotates, cooling fluid is supplied from the cooling source to each turbine vane segment outer band. The cooling fluid is channeled downstream through the outer band opening to the rotating blades. More specifically, the cooling fluid is supplied circumferentially around the rotor blade tips to facilitate reducing an operating temperature of the rotor blade tips and the shrouds surrounding the rotor blades. As a result, the turbine nozzle assembly facilitates reducing an operating temperature of the rotor assembly in a cost-effective and reliable manner.